Annular gas turbine housing component and a gas turbine comprising the component

ABSTRACT

The invention relates to an annular gas turbine housing component ( 15 ) comprising a plurality of circumferentially spaced first channels ( 16 ) for conveying a first gas flow in a first direction and a plurality of circumferentially spaced second channels ( 17 ) for conveying a second gas flow in a second direction across the first direction in a crossing position ( 18 ) of the first and second channels.

BACKGROUND AND SUMMARY

The present invention relates to an annular gas turbine housingcomponent. The invention is further directed to a gas turbine engine,and especially to an aircraft engine, comprising the component. Thus,the invention is especially, directed to a jet engine.

Jet engine is meant to include various types of engines, which admit airat relatively low velocity, heat it by combustion and shoot it out at amuch higher velocity. Accommodated within the term jet engine are, forexample, turbojet engines and turbofan engines. The invention will belowbe described for a turbofan engine, but may of course also be used forother engine types.

An aircraft gas turbine engine of the turbofan type generally comprisesa forward fan and booster compressor, a middle core engine, and an aftlow pressure power turbine. The core engine comprises a high pressurecompressor, a combustor and a high pressure turbine in a serialrelationship. The high pressure compressor and high pressure turbine ofthe core engine are interconnected by a high pressure shaft. Thehigh-pressure compressor is rotatably driven to compress air enteringthe core engine to a relatively high pressure. This high pressure air isthen mixed with fuel in the combustor and ignited to form a high energygas stream. The gas stream flows aft and passes through thehigh-pressure turbine, rotatably driving it and the high pressure shaftwhich, in turn, rotatably drives the high pressure compressor.

The gas stream leaving the high pressure turbine is expanded through asecond or low pressure turbine. The low pressure turbine rotatablydrives the fan and booster compressor via a low pressure shaft. The lowpressure shaft extends through the high pressure rotor. Most of thethrust produced is generated by the fan.

There is a desire for improved performance and fuel efficiency inturbofan and other turbine engines. Specific thrust may be increased andspecific fuel consumption may be decreased by increasing the cyclepressure ratio (CPR) of the engine. One known way of increasing the CPRbeyond today's level is to cool the incoming air between the compressorstages. The annular gas turbine housing component may be arrangedbetween the low pressure compressor and the high pressure compressor andconfigured to define the primary gas flow channel through the gasturbine engine.

U.S. Pat. No. 6,314,880 discloses a gas turbine engine provided with anintercooler connected to the core air flow and constructed to transportthe core air downstream of the low pressure compressor to the bypass airpassage and back to the core air passage upstream of the high pressurecompressor. More specifically, the air flow from the low pressurecompressor is conveyed in a loop via circumferentially spaced channelsprovided in the bypass air passage, which channels convey the core airin the opposite direction in relation to the bypass air flow, back tothe core air passage. A diffuser is provided for directing the core airflow to the circumferentially spaced channels in the bypass air passageand a collector is provided for collecting the air flow from thecircumferentially spaced channels in the bypass air passage beforeentering the high pressure compressor. The diffuser and the collectorare arranged in a crossing relationship so that the core air to theintercooler crosses the core air from the intercooler.

It is desirable to achieve a gas turbine housing component comprising awall structure, especially for application between compressor stages,which component is more rigid than prior art solutions and which createsconditions for low flow-losses. The improved rigidity creates conditionsfor an improved load-carrying ability. The component should further becost-efficient in production while maintaining or improving itsoperational characteristics.

According to an aspect of the invention, an annular gas turbine housingcomponent is provided comprising a plurality of circumferentially spacedfirst channels for conveying a first gas flow in a first directioncharacterized in that the component comprises a plurality ofcircumferentially spaced second channels for conveying a second gas flowin a second direction across the first direction in a crossing positionof the first and second channels.

Due to that the component comprises the plurality of circumferentiallyspaced second channels, which crosses the plurality of circumferentiallyspaced first channels, a rigid structure can be achieved.

Especially, a symmetric channel structure may be achieved with regard toa vertical plane along a center axis of the component and/or in acircumferential direction of the component. In order to achieve such arigid structure, according to one example, a plurality of the firstchannels alternate with a plurality of the second channels in thecircumferential direction.

According to a more specified example, every other channel in thecircumferential direction is a first channel and every other channel inthe circumferential direction is a second channel at least along aportion of the circumference. Preferably, every other channel in thecircumferential direction is a first channel and every other channel inthe circumferential direction is a second channel at least along asubstantial part of the circumference. More preferably, every otherchannel in the circumferential direction is a first channel and everyother channel in the circumferential direction is a second channel alongthe complete circumference.

Further, by arranging the plurality of circumferentially spaced secondchannels in a crossing relationship with the plurality ofcircumferentially spaced first channels, arrangements of sub-systems maybe facilitated. For example, the radially spaced channels createsconditions for radially extending elements, such as power take offshafts, oil conduits etc.

According to a further example embodiment, the component is configuredfor a channel loop connecting said first channels to said secondchannels so that a gas flow from said first channels subsequently willflow through said second channels in gas turbine operation.

Such a design creates conditions for cooling the gas flow downstream ofthe first channels and upstream of the second channels. Preferably, aheat exchanger is positioned in said loop for cooling the gas flow.

According to an example embodiment, at least a plurality of the firstchannels are configured so that there is an increase in their radialextension from a position upstream the crossing position in a gas flowdirection to the crossing position. This diverging design of the firstchannels creates conditions for limiting a flow path area decrease,which would otherwise, be necessitated due to the crossing of thechannels. In other words, a flow speed increase, which would otherwisebe necessitated, is limited. Such an arrangement is for exampledesirable in case there is a heat exchanger provided in the loopaccording to the last-mentioned embodiment since the flow losses in theheat exchanger are correspondingly limited.

According to a further example embodiment, said plurality of firstchannels are configured so that a flow path area in said crossingposition is at least as large as in a position upstream the crossingposition in a gas flow direction. The “flow path area” is here definedas the total flow path area in all the first channels. Further, the“upstream position” is located at an exit of the upstream compressorstage in the gas turbine or a position inbetween the upstream compressorstage exit and the crossing position. A flow speed decrease can therebybe achieved. Such an arrangement is for example desirable in case thereis a heat exchanger provided in the loop according to the last-mentionedembodiment since the flow losses in the heat exchanger arecorrespondingly limited.

According to a further example embodiment, said plurality of firstchannels are configured so that a flow path area in a position upstreamthe crossing position in a gas flow direction is substantially the sameas in said crossing position. Thus, the speed of the gas flow issubstantially the same in the upstream position and the crossingposition.

Other advantageous features and functions of various embodiments of theinvention are set forth in the following description and in thedependent claims.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be explained below, with reference to the embodimentshown on the appended drawings, wherein

FIG. 1 is a schematic side view of an aircraft engine cut along a planein parallel with the rotational axis of the engine,

FIGS. 2 and 3 are two schematic views of the flow channels in anintermediate housing component from FIG. 1,

FIG. 4 is a schematic, perspective view of a crossing point of the flowchannels in FIGS. 2 and 3, and

FIGS. 5 and 6 shows two different embodiments of the flow channels inthe intermediate housing component.

DETAILED DESCRIPTION

The invention will below be described for a turbofan gas turbineaircraft engine 1, which in FIG. 1 is circumscribed about an enginelongitudinal central axis 2. The engine 1 comprises an outer casing ornacelle 3, an inner casing 4 (rotor) and an intermediate casing 5 whichis concentric to the first two casings and divides the gap between theminto an inner primary gas channel 6 for the compression of air and asecondary channel 7 in which the engine bypass air flows. The casingsare in turn made up of a plurality of components in the axial directionof the engine. Thus, each of the gas channels 6,7 is annular in a crosssection perpendicular to the engine longitudinal central axis 2.

The engine 1 comprises a fan 8 which receives ambient air 9, a boosteror low pressure compressor (LPC) 10 and a high pressure compressor (HPC)11 arranged in the primary gas channel 6, a combustor 12 which mixesfuel with the air pressurized by the high pressure compressor 11 forgenerating combustion gases which flow downstream through a highpressure turbine (HPT) 13 and a low pressure turbine (LPT) 14 from whichthe combustion gases are discharged from the engine.

A high pressure shaft joins the high pressure turbine 13 to the highpressure compressor 11 to substantially form a high pressure rotor. Alow pressure shaft joins the low pressure turbine 14 to the low pressurecompressor 10 to substantially form a low pressure rotor. The lowpressure shaft is at least in part rotatably disposed co-axially withand radially inwardly of the high pressure rotor.

An annular gas turbine housing component 15 is positioned between thelow pressure compressor 10 and the high pressure compressor 11. Thecomponent 15 will now be further described with reference to FIGS. 2-5.The arrows indicate the flow direction. The component 15 comprises aplurality of circumferentially spaced first channels 16 for conveying afirst gas flow in a first direction and a plurality of circumferentiallyspaced second channels 17 for conveying a second gas flow in a seconddirection across the first direction in a crossing position 18 of thefirst and second channels. A gas flow from the low pressure compressor10 will enter the first channels 16 and the gas flow from the secondchannels 17 will enter the high pressure compressor 11 in operation.

More specifically, a plurality of the first channels 16 alternate with aplurality of the second channels 17 in the circumferential direction.Especially, every other channel in the circumferential direction is afirst channel 16 and every other channel in the circumferentialdirection is a second channel 17 along the complete circumference.

The component 15 comprises a wall structure 19, which comprises aplurality of circumferentially spaced radial walls 20,120,220, whichdefine said first channels 16 and second channels 17, see FIG. 3. Thewall structure 19 is configured so that each of said radial walls20,120,220 define one of said first channels 16 on one side and one ofsaid second channels 17 on the other side in the circumferentialdirection at least along a portion of the circumference and preferablyalong the complete circumference. The wall structure is configured to beload-carrying.

At least a plurality of the first channels 16 are configured so thatthere is an increase 21 in their radial extension from a position 22upstream the crossing position 18 in a gas flow direction to thecrossing position 18. Each of said plurality of first channels 16 isconfigured so that a flow path area in said upstream position issubstantially the same as in said crossing position. Alternatively, theflow path area in said upstream position may be smaller than in saidcrossing position. In any case, the gas flow path area is notdrastically decreased in said crossing position in relation to saidupstream position.

The component 15 comprises a circumferentially substantially continuousgas flow channel 23, which is divided into said plurality ofcircumferentially spaced first channels 16. An axial position of saidchannel division is about at the same axial position as the upstreamposition. In the shown embodiment, the upstream position 22 is definedby an axial position of said channel division.

The first channels 16 are configured for conveying the first gas flowsubstantially in an axial direction of the component 15 past thecrossing point 18. More specifically, a center line 24 of the continuousgas flow channel 23 in a radial direction in the upstream position 22 ispositioned on substantially the same radial position as a center line 25of said plurality of circumferentially spaced first channels 16 in saidcrossing position 18. Further, the plurality of second channels 17 areconfigured for conveying the second gas flow substantially in a radialdirection of the component 15 past the crossing point 18. Accordingly,the first channels 16 are directed substantially perpendicular to thesecond channels 17 at the crossing point 18.

The component 15 creates conditions for a channel loop 26 connectingsaid first channels 16 to said second channels 17 so that a gas flowfrom said first channels 16 subsequently will flow through said secondchannels 17 in gas turbine operation. More specifically, a heatexchanger 27 is positioned in said loop for cooling the gas flow. Theheat exchanger preferably comprises a cooling structure 28 withplurality of spaced tubes 29 for conveying the gas flow from the firstchannels 16 to the second channels 17. A plurality of cooling fins 30are arranged between adjacent' cooling tubes 29 for improving thecooling performance.

In order to limit flow losses in the heat exchanger 27, the gas flowarea should be substantially the same or may even be increased from anoutlet of the upstream compressor (low pressure compressor) 10 to aninlet of the heat exchanger 27. Thus, the flow path area may beincreased from the outlet of the upstream compressor to the crossingposition 18 and preferably also downstream of the crossing position 18towards the heat exchanger 27. Further, the gas flow area should besubstantially the same or may even be decreased from an outlet of theheat exchanger 27 to an inlet of the downstream compressor (highpressure compressor) 11.

The heat exchanger 27 preferably comprises a plurality of exchangeablecooling modules. According to a first example, a plurality of suchcooling modules is arranged as sectors, which together form an annularcooling structure. According to an alternative, or complement to thefirst example, a plurality of such cooling modules is arranged next toeach others in an axial direction of the gas turbine engine.

According to a further alternative, the heat exchanger may comprise aplurality of heat exchanging element arranged in series in said loop 26.

According to the embodiment shown in FIG. 1, the heat exchanger 27 ispositioned radially interior of the intermediate casing 5. Morespecifically, the bypass channel is divided and the heat exchanger 27 ispositioned in a secondary by-pass channel. In other words, the heatexchanger 27 is positioned radially interior of the main bypass passage7. This position of the heat exchanger is possible due to the fact thata downstream end of the low pressure compressor 10 is positioned on asubstantially larger radial distance than an upstream end of the highpressure compressor 11, see FIG. 1.

The design of the second channels 17 in FIG. 4 should only be regardedas a schematic example showing the function of the invention. In casethey are positioned in a flow channel (such as the bypass channel), theyare preferably aerodynamically shaped for causing as small disturbanceas possible to the passing, bypass flow.

The inventive component 15 is configured for a core engine, i.e anengine with or without a bypass channel.

In the application shown in FIG. 1, the component is arranged so thatuncombusted air is conveyed via the second channels 17 during operation.Further, the component is arranged so that the second gas flow in thesecond channels flows from a radially outer position to a radially innerposition in relation to the crossing point during operation. Further,the gas turbine comprises a channel loop 26 connecting said firstchannels 16 to said second channels 17 so that a gas flow from saidfirst channels subsequently will flow through said second channels ingas turbine operation. Further, the gas turbine is configured to coolthe gas flow in said loop 26 during operation.

A second embodiment of the component is indicated in FIG. 6. The firstchannels 116 are directed at an oblique angle with regard to the axialdirection. Thus, the direction of the first channels 116 has a componentin the radial direction while the second channels are directedsubstantially in the radial direction. In other words, the firstchannels 116 cross the second channels 117 in an angle substantially,different from 90°.

According to one example, the loop 26 connects two adjacent channels16,17 in the circumferential direction. In other words, a gas flow inone of said first channels' 16 returns in a neighbouring second channel17. According to an alternative, the loop may be arranged so that thegas flow in one of said first channels 16 may be returned in a secondchannel 17 at any circumferential distance from the first channel 16.According to a further alternative, the component may be configured sothat the gas flow from a plurality of said first channels 16 is joinedbefore it is returned in one or a plurality of said second channels 17.

The invention is not in any way limited to the above describedembodiments, instead a number of alternatives and modifications arepossible without departing from the scope of the following claims.

For example, the heat exchanger may be located in a secondary gas flowchannel (by-pass channel) or in engine configurations without a by-passchannel, on the exterior of the engine housing. Further, the design ofthe heat exchanger should only be regarded as an example.

Further, the direction of the first channels and the direction of thesecond channels in the crossing point may vary in differentapplications. Further, the direction of different first channels and/orsecond channels may vary along the circumference of the component.

Further, according to an alternative to the wall structure embodimentshown in FIG. 2-5, the first and second channels may be formed by tubescrossing each other. Further, the component 15 may be formed by joininga plurality of sub-components. For example, the first channels may beintegrated in a first subcomponent and the second channels may beintegrated in a second sub-component.

According to an alternative to the shown embodiments, the channeldivision is located upstreams (or downstreams) of the position where theradial extension of the first channels is increased.

1. An annular gas turbine housing component comprising a plurality ofcircumferentially spaced first channels for conveying a first gas flowin a first direction, and a plurality of circumferentially spaced secondchannels for conveying a second gas flow in a second direction acrossthe first direction in a crossing position of the first and secondchannels.
 2. A gas turbine housing component according to claim 1,wherein a plurality of the first channels alternate with a plurality ofthe second channels in the circumferential direction.
 3. A gas turbinehousing component according to claim 1, wherein every other channel inthe circumferential direction is a first channel and every other channelin the circumferential direction is a second channel at least along aportion of the circumference.
 4. A gas turbine housing componentaccording to claim 1, wherein the component comprises a wall structure,which comprises a plurality of circumferentially spaced radial walls,which define the first channels and second channels.
 5. A gas turbinehousing component according to claim 4, wherein the wall structure isconfigured so that each of the radial walls define one of the firstchannels on one side and one of the second channels on the other side inthe circumferential direction at least along a portion of thecircumference.
 6. A gas turbine housing component according to claim 1,wherein at least a plurality of the first channels are configured sothat there is an increase in their radial extension from a positionupstream the crossing position in a gas flow direction to the crossingposition.
 7. A gas turbine housing component according to claim 1wherein the plurality of first channels are configured so that a flowpath area in the crossing position is at least as large as in a positionupstream the crossing position in a gas flow direction.
 8. A gas turbinehousing component according to claim 1, wherein the plurality of firstchannels are configured so that a flow path area in a position upstreamthe crossing position in a gas flow direction is substantially the sameas in the crossing position.
 9. A gas turbine housing componentaccording to claim 1 wherein the component comprises a circumferentiallysubstantially continuous gas flow channel, which is divided into theplurality of circumferentially spaced first channels.
 10. A gas turbinehousing component according to claim 9, wherein an axial position of thechannel division is about at the same axial position as the upstreamposition.
 11. A gas turbine housing component according to claim 10,wherein a center line of the continuous gas flow channel in a radialdirection in the upstream position is positioned on substantially thesame radial position as a center line of the plurality ofcircumferentially spaced first channels in the crossing position.
 12. Agas turbine housing component according to claim 1, characterized inthat the first channels are configured for conveying the first gas flowsubstantially in an axial direction (2) of the component.
 13. A gasturbine housing component according to claim 1, wherein the plurality ofsecond channels are configured for conveying the second gas flowsubstantially in a radial direction of the component.
 14. A gas turbinehousing component according to claim 1, wherein the component isconfigured for a channel loop connecting the first channels to thesecond channels so that a gas flow from the first channels subsequentlywill flow through the second channels in gas turbine operation.
 15. Agas turbine housing component according to claim 14, wherein heatexchanger is positioned in the loop for cooling the gas flow.
 16. A gasturbine housing component according to claim 1, wherein the first gasflow channel defines a primary gas flow channel.
 17. A gas turbinewherein it comprises a gas turbine housing component according toclaim
 1. 18. A gas turbine according to claim 17, wherein the componentis arranged so that uncombusted air is conveyed via the second channelsduring operation.
 19. A gas turbine according to claim 17, wherein thecomponent is arranged so that the second gas flow in the second channelsflows from a radially outer position to a radially inner position inrelation to the crossing point during operation.
 20. A gas turbineaccording to claim 17, wherein the gas turbine comprises a channel loopconnecting the first channels to the second channels so that a gas flowfrom the first channels subsequently will flow through the secondchannels in gas turbine operation.
 21. A gas turbine according to claim20, wherein the gas turbine is configured to cool the gas flow in theloop during operation.
 22. A gas turbine according to claim 17, whereinit comprises at least two axially spaced compressor stages and that thegas turbine housing component is positioned between the compressorstages so that a gas flow from a first compressor stage will enter thefirst channels and the gas flow from the second channels will enter asecond compressor stage in operation.
 23. A gas turbine according toclaim 22 wherein a downstream end of the first compressor stage ispositioned on a substantially larger radial distance than an upstreamend of the second compressor stage.
 24. An annular gas turbine housingcomponent comprising a plurality of circumferentially spaced firstchannels for conveying a first gas flow in a first direction, whereinthe component comprises a plurality of circumferentially spaced secondchannels for conveying a second gas flow in a second direction acrossthe first direction in a crossing position of the first and secondchannels.
 25. A gas turbine housing component according to claim 24,wherein a plurality of the first channels alternate with a plurality ofthe second channels in the circumferential direction.
 26. A gas turbinehousing component according to claim 24, wherein every other channel inthe circumferential direction is a first channel and every other channelin the circumferential direction is a second channel at least along aportion of the circumference.
 27. A gas turbine housing componentaccording to claim 24, wherein the component comprises a wall structure,which comprises a plurality of circumferentially spaced radial walls,which define the first channels and second channels.
 28. A gas turbinehousing component according to claim 27, wherein the wall structure isconfigured so that each of the radial walls define one of the firstchannels on one side and one of the second channels on the other side inthe circumferential direction at least along a portion of thecircumference.
 29. A gas turbine housing component according to claim24, wherein at least a plurality of the first channels are configured sothat there is an increase in their radial extension from a positionupstream the crossing position in a gas flow direction to the crossingposition.
 30. A gas turbine housing component according to claim 24,wherein the plurality of first channels are configured so that a flowpath area in the crossing position is at least as large as in a positionupstream the crossing position in a gas flow direction.
 31. A gasturbine housing component according to claim 24, wherein the pluralityof first channels are configured so that a flow path area in a positionupstream the crossing position in a gas flow direction is substantiallythe same as in the crossing position.
 32. A gas turbine housingcomponent according to claim 24, wherein the component comprises acircumferentially substantially continuous gas flow channel, which isdivided into the plurality of circumferentially spaced first channels.33. A gas turbine housing component according to claim 32, wherein anaxial position of the channel division is about at the same axialposition as the upstream position.
 34. A gas turbine housing componentaccording to claim 33, wherein a center line of the continuous gas flowchannel in a radial direction in the upstream position is positioned onsubstantially the same radial position as a center line of the pluralityof circumferentially spaced first channels in the crossing position. 35.A gas turbine housing component according to claim 24, wherein the firstchannels are configured for conveying the first gas flow substantiallyin an axial direction of the component.
 36. A gas turbine housingcomponent according to claim 24, wherein the plurality of secondchannels are configured for conveying the second gas flow substantiallyin a radial direction of the component.
 37. A gas turbine housingcomponent according to claim 24, wherein the component is configured fora channel loop connecting the first channels to the second channels sothat a gas flow from the first channels subsequently will flow throughthe second channels in gas turbine operation.
 38. A gas turbine housingcomponent according to claim 37, wherein heat exchanger is positioned insaid the loop for cooling the gas flow.
 39. A gas turbine housingcomponent according to claim 24, wherein the first gas flow channeldefines a primary gas flow channel.
 40. A gas turbine wherein itcomprises a gas turbine housing component according to claim
 24. 41. Agas turbine according to claim 40, wherein the component is arranged sothat uncombusted air is conveyed via the second channels duringoperation.
 42. A gas turbine according to claim 40, wherein thecomponent is arranged so that the second gas flow in the second channelsflows from a radially outer position to a radially inner position inrelation to the crossing point during operation.
 43. A gas turbineaccording to claim 40, wherein the gas turbine comprises a channel loopconnecting the first channels to the second channels so that a gas flowfrom the first channels subsequently will flow through the secondchannels in gas turbine operation.
 44. A gas turbine according to claim43, wherein the gas turbine is configured to cool the gas flow in theloop during operation.
 45. A gas turbine according to claim 40, whereinit comprises at least two axially spaced compressor stages and that thegas turbine housing component is positioned between the compressorstages so that a gas flow from a first compressor stage will enter thefirst channels and the gas flow from the second channels will enter asecond compressor stage in operation.
 46. A gas turbine according toclaim 45 wherein a downstream end of the first compressor stage ispositioned on a substantially larger radial distance than an upstreamend of the second compressor stage.